A solid-fuel air turbine rocket engine

Abstract

This invention relates to the field of engine technology and proposes a solid-fuel air turbine rocket engine, comprising a casing, a gas generator, a turbine, and an axial compressor. A central cone is fixedly disposed within the casing; the gas generator is fixedly disposed within the casing; the turbine includes turbine rotor blades and turbine stator blades; the axial compressor includes compressor stator blades and compressor rotor blades, wherein the compressor rotor blades are located outside the turbine stator blades and are fixedly connected to the turbine rotor blades, and the compressor rotor blades are rotatably connected to the compressor stator blades. This invention, by placing the compressor rotor blades around the turbine rotor blades and fixing them together, positions the gas generator at the front end of the turbine and compressor, increasing the installation space for the gas generator and allowing more solid fuel to be loaded into it, thereby improving the engine's range.

Description

Technical Field

[0001] This invention relates to the field of engine technology, and more particularly to a solid-fuel air turbine rocket engine. Background Technology

[0002] With the development of human aerospace technology, various types of aircraft have emerged, and the requirements for propulsion systems are constantly increasing. Air-turbo rocket engines (ATR engines) have become one of the alternative power solutions for reusable near-space vehicles due to their wide speed range and zero-speed start-up capabilities. Furthermore, solid-fuel air-turbo rocket engines (SP-ATR engines) can serve as missile propulsion systems, meeting the requirements for high-thrust acceleration, low-thrust cruise, and terminal acceleration.

[0003] Patent publication number CN111271192A discloses an air turbine rocket engine based on pulse detonation, mainly composed of an air intake, compressor, gas generator, turbine, mixing and anti-reverse transmission structure, pulse detonation combustion chamber, and nozzle. The engine contains a gas generator carrying oxidizer and fuel. The fuel-rich gas generated by the gas generator drives the turbine to do work, and the turbine drives the compressor. Air entering from the air intake is compressed by the compressor and then mixed evenly with the fuel-rich gas after it has been powered by the turbine in the mixing unit before entering the pulse detonation combustion chamber for detonation combustion. The high-temperature, high-pressure gas generated by detonation combustion is discharged through the nozzle to generate thrust.

[0004] In the above-mentioned technical solution, the compressor and turbine are coaxially arranged, with the gas generator located between the compressor and the turbine. When the engine is working, the fuel-rich gas generated by the gas generator drives the turbine. The turbine drives the compressor through the central shaft to compress the air drawn in from the intake. The compressed air and fuel-rich gas are mixed and enter the combustion chamber for combustion. Finally, the gas is ejected through the nozzle to generate thrust. However, with this solution, the space of the gas generator inside the engine is limited, resulting in a smaller amount of solid fuel, which is difficult to meet the needs of rockets for long-term flight. Summary of the Invention

[0005] In view of this, the present invention proposes a solid fuel air turbine rocket engine, in which the compressor moving blades are arranged on the periphery of the turbine moving blades and the two are fixedly connected, which can increase the installation space of the gas generator and increase the amount of solid fuel propellant.

[0006] The technical solution of this invention is implemented as follows: This invention provides a solid-fuel air turbine rocket engine, comprising a casing, a gas generator, a turbine, and an axial-flow compressor, wherein,

[0007] A central cone is fixedly installed inside the housing;

[0008] The gas generator is fixedly installed inside the housing, and the gas generator has an opening at one end near the central cone for injecting fuel-rich gas into the housing.

[0009] The turbine includes a turbine moving blade and a turbine stationary blade, wherein the turbine moving blade is rotatably disposed on the outside of the central cone; the turbine stationary blade is fixedly disposed on the inside of the opening and is rotatably connected to the turbine moving blade.

[0010] The axial compressor includes compressor stationary blades and compressor moving blades, wherein the compressor stationary blades are fixedly disposed within the housing; the compressor moving blades are located outside the turbine stationary blades and are fixedly connected to the turbine moving blades, and the compressor moving blades are rotatably connected to the compressor stationary blades.

[0011] Based on the above technical solutions, preferably, both the turbine moving blade and the turbine stationary blade are provided with multiple stages, and the multiple stages of the turbine moving blade and the multiple stages of the turbine stationary blade are arranged at intervals.

[0012] Both the compressor stationary blades and the compressor moving blades are provided in multiple stages, with the multiple stages of compressor stationary blades and the multiple stages of compressor moving blades arranged at intervals, and the multiple stages of compressor moving blades are fixedly connected to each other.

[0013] More preferably, the turbine blade with the opening at one end near the gas generator and the compressor blade with the opening at one end near the gas generator are fixedly connected by a connector.

[0014] More preferably, the connector is an annular plate structure that seals the gap between the fixedly connected turbine blade and the compressor blade.

[0015] Based on the above technical solutions, preferably, a cone is fixedly provided at the end of the gas generator away from the opening.

[0016] More preferably, the gas generator is coaxially arranged with the housing, and an air inlet is provided between the gas generator and the housing.

[0017] More preferably, the cross-sectional area of ​​the end of the air intake duct away from the central cone first decreases and then increases, and the end of the air intake duct near the central cone is a columnar structure with a circular cross-section.

[0018] More preferably, it also includes a mixer, which is fixedly disposed within the housing and located on the side of the central cone away from the air intake.

[0019] More preferably, a nozzle is fixedly provided at the end of the housing away from the air intake, and the position between the mixer and the nozzle inside the housing is the combustion chamber;

[0020] The mixer has a first flow channel and multiple second flow channels. The first flow channel is located at the center of the mixer. One end of the first flow channel is connected to the combustion chamber, and the other end is connected to the opening through the turbine. Multiple second flow channels are arranged circumferentially around the center line of the first flow channel. One end of the second flow channel is connected to the combustion chamber, and the other end is connected to the intake through the axial compressor.

[0021] More preferably, the nozzle, the turbine moving blade, the turbine stationary blade, the compressor stationary blade, and the compressor moving blade are coaxially arranged with the housing.

[0022] The solid-fuel air turbine rocket engine of the present invention has the following advantages over the prior art:

[0023] (1) By setting the compressor moving blades on the periphery of the turbine moving blades and fixing the two together, the gas generator is located at the front end of the turbine and compressor, which increases the installation space of the gas generator and allows more solid fuel to be filled in the gas generator, thereby improving the range performance of this engine.

[0024] (2) By setting a connector to fix the last stage compressor blade to the last stage turbine blade, not only can the installation restrictions of the turbine and compressor be reduced, but the risk of fuel leakage can also be reduced.

[0025] (3) The fuel-rich gas generated by the gas generator is directly delivered to the turbine, eliminating the need for the pipeline delivery structure used in traditional solid fuel air turbine rocket engines, making the structure of this engine more streamlined. Attached Figure Description

[0026] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0027] Figure 1 This is a cross-sectional view of a solid fuel air turbine rocket engine according to the present invention;

[0028] Figure 2 for Figure 1 Enlarged view of point A in the middle;

[0029] Figure 3 This is a cross-sectional view of a connecting component in a solid fuel air turbine rocket engine according to the present invention;

[0030] Figure 4 This is a cross-sectional view of the turbine stator blade in a solid fuel air turbine rocket engine according to the present invention.

[0031] The components are: 1. Shell; 11. Central cone; 12. Nozzle; 101. Inlet; 102. Combustion chamber; 2. Gas generator; 21. Cone; 201. Opening; 3. Turbine; 31. Turbine moving blade; 311. Connecting part; 32. Turbine stationary blade; 4. Axial compressor; 41. Compressor stationary blade; 42. Compressor moving blade; 5. Mixer; 501. First flow channel; 502. Second flow channel. Detailed Implementation

[0032] The technical solutions of this invention will be clearly and completely described below with reference to specific embodiments. Obviously, the described embodiments are only a part of the embodiments of this invention, and not all of them. Based on the embodiments of this invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this invention.

[0033] like Figure 1-4 As shown, a solid fuel air turbine rocket engine of the present invention includes a housing 1, a gas generator 2, a turbine 3, an axial compressor 4, and a mixer 5.

[0034] Among them, the housing 1 is the outer shell structure of the engine, which is usually cylindrical and is used to support and protect other components of the engine. A central cone 11 is fixedly installed inside the housing 1 to support the turbine 3.

[0035] Gas generator 2 is used to store solid fuel for the engine. Gas generator 2 is fixedly installed inside housing 1. Gas generator 2 has an opening 201 at one end near the central cone 11, such as... Figure 1 As shown, the gas generator 2 is preferably located at the left end of the housing 1, while the opening 201 is opened at the right end of the gas generator 2. After the solid fuel in the gas generator 2 is ignited, the fuel-rich gas is injected into the interior of the housing 1.

[0036] The turbine 3 includes a turbine moving blade 31 and a turbine stationary blade 32. The turbine moving blade 31 is fixedly disposed on the outer side of the central cone 11. The turbine stationary blade 32 is fixedly disposed on the inner side of the opening 201 and is rotatably connected to the turbine moving blade 31. When the gas generator 2 sprays rich gas, the rotatable turbine moving blade 31 is driven to rotate by the rotational engagement of the stationary turbine stationary blade 32 (this rotational engagement is prior art), which converts the enthalpy of the airflow into the kinetic energy and mechanical work of the rotating part and delivers the rich gas into the mixer 5.

[0037] The axial compressor 4 provides high-pressure air to the engine. The axial compressor 4 includes a compressor stator vane 41 and a compressor moving vane 42. The compressor stator vane 41 is fixedly installed inside the housing 1. The compressor moving vane 42 is located outside the turbine stator vane 32 and is fixedly connected to the turbine moving vane 31. The compressor moving vane 42 and the compressor stator vane 41 are rotatably connected. Figure 1 As shown, the turbine stationary blade 32 is fixedly installed inside the gas generator 2. The compressor stationary blade 41 and the compressor moving blade 42 are both located outside the gas generator 2 at positions that coincide with the turbine stationary blade 32 and the turbine moving blade 31. When the fuel-rich gas drives the turbine moving blade 31 to rotate, it can drive the compressor moving blade 42, which is fixedly connected to the turbine moving blade 31, to rotate. Then, by utilizing the rotational cooperation between the compressor moving blade 42 and the compressor stationary blade 41 (this rotational cooperation is also existing technology), the incoming air at the inlet of the axial compressor 4 is compressed.

[0038] In a preferred embodiment, in order to improve the driving effect of the turbine 3 and the air compression effect of the axial compressor 4, the turbine moving blade 31, turbine stationary blade 32, compressor stationary blade 41 and compressor moving blade 42 can all be set to multiple stages, i.e. multiple turns, with the multi-stage turbine moving blade 31 and the multi-stage turbine stationary blade 32 arranged alternately, the multi-stage compressor stationary blade 41 and the multi-stage compressor moving blade 42 arranged alternately, and the multi-stage compressor moving blade 42 fixedly connected to each other.

[0039] Regarding the fixed connection between the turbine blade 31 and the compressor blade 42, it is preferable to fix the turbine blade 31, which is located near the end of the gas generator 2 with the opening 201, and the compressor blade 42, which is located near the end of the gas generator 2 with the opening 201, together with the turbine blade 311. Figure 1 and Figure 2As shown, the left end of the rocket engine is the front section. The leftmost first-stage turbine blade 31 (i.e., the leftmost ring) is the first-stage turbine blade 31, and the rightmost first-stage turbine blade 31 is the last-stage turbine blade 31. The leftmost first-stage compressor blade 42 is the first-stage compressor blade 42, and the rightmost first-stage compressor blade 42 is the last-stage compressor blade 42. By using the connector 311 to fix the last-stage turbine blade 31 and the last-stage compressor blade 42, the adhesion between the multi-stage turbine blades 31 and the multi-stage compressor blades 42 can be reduced, simplifying the combined structure of the two, thereby reducing the installation restrictions of the turbine 3 and the axial compressor 4.

[0040] The connector 311 is preferably configured as an annular plate structure. The connector 311 is used to seal the gap between the fixedly connected turbine blade 31 and compressor blade 42. That is, the connector 311 covers the gap between the last stage turbine blade 31 and the last stage compressor blade 42. This allows the fuel-rich gas to be transported along the channels between each turbine blade 31 and each turbine stator blade 32 (the channels between multiple turbine blades 31 in the same stage, the channels between multiple turbine stator blades 32 in the same stage, and the channels between adjacent turbine blades 31 and turbine stator blades 32) without leaking out from the gap between the turbine blades 31 and the compressor blade 42, thereby reducing the risk of fuel leakage.

[0041] To ensure uniform airflow into the housing 1, it is preferable to align the gas generator 2 coaxially with the housing 1. Figure 1 As shown, an air intake duct 101 is provided between the gas generator 2 and the housing 1. Air enters the engine through the air intake duct 101 and is compressed by the axial compressor 4 and delivered to the mixer 5. Specifically, a cone 21 is fixedly provided at the end of the gas generator 2 away from the opening 201, such as... Figure 1 As shown, by utilizing the cooperation of the cone 21, the housing 1 and the gas generator 2, the cross-sectional area of ​​the intake duct 101 at the end away from the central cone 11 can be reduced first and then increased. The end of the intake duct 101 near the central cone 11 is a columnar structure with a circular cross-section. With this structure, when air enters the engine from left to right, the intake duct 101 first contracts and then expands, and then flows evenly and smoothly.

[0042] The mixer 5 is used to mix air and fuel-rich gas. The mixer 5 is fixedly installed inside the housing 1 and located on the side of the central cone 11 away from the intake duct 101. The nozzle 12 is fixedly installed at the end of the housing 1 away from the intake duct 101. The position between the mixer 5 and the nozzle 12 inside the housing 1 is the combustion chamber 102. The air and fuel-rich gas are mixed and burned in the combustion chamber 102. The gas after combustion is ejected to the outside through the nozzle 12 to provide thrust for the engine.

[0043] To better achieve the fusion of air and fuel-rich gas, it is preferable to open a first flow channel 501 and multiple second flow channels 502 in the mixer 5. The first flow channel 501 is located at the center of the mixer 5. One end of the first flow channel 501 is connected to the combustion chamber 102, and the other end is connected to the opening 201 through the turbine 3, so that the fuel-rich gas can quickly enter the combustion chamber 102. Multiple second flow channels 502 are arranged circumferentially around the center line of the first flow channel 501. One end of the second flow channel 502 is connected to the combustion chamber 102, and the other end is connected to the intake duct 101 through the axial compressor 4, so that the air and fuel-rich gas can be fully fused.

[0044] Furthermore, it is preferable to arrange the nozzle 12, turbine moving blade 31, turbine stationary blade 32, compressor stationary blade 41, and compressor moving blade 42 coaxially with the casing 1 to ensure the uniformity of air and fuel-rich gas flow and the pressure balance of various parts of the rocket engine.

[0045] The working principle of a solid-fuel air turbine rocket engine according to the present invention is as follows:

[0046] like Figure 1 As shown, the gas generator 2 sprays the solid fuel inside into a rich gas mixture to the right, which drives the turbine blade 31 to rotate and enters the combustion chamber 102 along the first flow channel 501. The turbine blade 31 can drive the compressor blade 42, which is fixedly connected to it, to rotate, thereby making the axial compressor 4 work. At the same time, air enters the intake duct 101 through the inlet, is decelerated and pressurized, and is then compressed by the axial compressor 4 at the outlet of the intake duct 101. The air then enters the combustion chamber 102 along the intake duct 101, the axial compressor 4, and the second flow channel 502, where it mixes and burns with the rich gas mixture. The burned gas is then sprayed out along the nozzle 12, thereby generating thrust for the engine.

[0047] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A solid-fuel air turbine rocket engine, characterized in that: It includes a casing (1), a gas generator (2), a turbine (3), and an axial compressor (4), wherein, A central cone (11) is fixedly installed inside the shell (1); The gas generator (2) is fixedly installed inside the housing (1). The gas generator (2) has an opening (201) at one end near the central cone (11) for injecting fuel-rich gas into the housing (1). The turbine (3) includes a turbine moving blade (31) and a turbine stationary blade (32), wherein the turbine moving blade (31) is rotatably disposed on the outside of the central cone (11); the turbine stationary blade (32) is fixedly disposed on the inside of the opening (201) and is rotatably connected to the turbine moving blade (31). The axial compressor (4) includes a compressor stationary blade (41) and a compressor moving blade (42). The compressor stationary blade (41) is fixedly disposed inside the housing (1). The compressor moving blade (42) is located outside the turbine stationary blade (32) and is fixedly connected to the turbine moving blade (31). The compressor moving blade (42) is rotatably connected to the compressor stationary blade (41).

2. The solid-fuel air turbine rocket engine as described in claim 1, characterized in that: Both the turbine moving blade (31) and the turbine stationary blade (32) are provided with multiple stages, and the multiple stages of the turbine moving blade (31) and the multiple stages of the turbine stationary blade (32) are arranged at intervals; Both the compressor stationary blade (41) and the compressor moving blade (42) are provided with multiple stages, and the multiple stages of the compressor stationary blade (41) and the multiple stages of the compressor moving blade (42) are arranged at intervals, and the multiple stages of the compressor moving blade (42) are fixedly connected to each other.

3. A solid-fuel air turbine rocket engine as described in claim 2, characterized in that: The turbine blade (31) located near the end of the opening (201) of the gas generator (2) is fixedly connected to the compressor blade (42) located near the end of the opening (201) of the gas generator (2) by a connector (311).

4. A solid-fuel air turbine rocket engine as described in claim 3, characterized in that: The connector (311) is an annular plate structure that seals the gap between the turbine blade (31) and the compressor blade (42) that are fixedly connected.

5. A solid-fuel air turbine rocket engine as described in claim 1, characterized in that: The gas generator (2) has a cone (21) fixedly provided at the end away from the opening (201).

6. A solid-fuel air turbine rocket engine as described in claim 5, characterized in that: The gas generator (2) is coaxially arranged with the housing (1), and an air inlet (101) is provided between the gas generator (2) and the housing (1).

7. A solid-fuel air turbine rocket engine as described in claim 6, characterized in that: The cross-sectional area of ​​the air intake (101) at the end away from the central cone (11) first decreases and then increases, and the end of the air intake (101) near the central cone (11) is a columnar structure with a circular cross-section.

8. A solid-fuel air turbine rocket engine as described in claim 6 or 7, characterized in that: It also includes a mixer (5), which is fixedly disposed inside the housing (1) and located on the side of the central cone (11) away from the air intake (101).

9. A solid-fuel air turbine rocket engine as described in claim 8, characterized in that: A nozzle (12) is fixedly provided at one end of the housing (1) away from the air intake (101), and the position between the mixer (5) and the nozzle (12) inside the housing (1) is the combustion chamber (102); The mixer (5) has a first flow channel (501) and a plurality of second flow channels (502). The first flow channel (501) is located at the center of the mixer (5). One end of the first flow channel (501) is connected to the combustion chamber (102), and the other end is connected to the opening (201) through the turbine (3). The plurality of second flow channels (502) are arranged circumferentially around the center line of the first flow channel (501). One end of the second flow channel (502) is connected to the combustion chamber (102), and the other end is connected to the intake duct (101) through the axial compressor (4).

10. A solid-fuel air turbine rocket engine as described in claim 9, characterized in that: The nozzle (12), the turbine moving blade (31), the turbine stationary blade (32), the compressor stationary blade (41), the compressor moving blade (42) are coaxially arranged with the housing (1).

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